Unducted thrust producing system

ABSTRACT

A unshrouded vane assembly for an unducted propulsion system includes a plurality of vanes which have non-uniform characteristics configured to generate a desired vane exit swirl angle.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application claims the benefit of the earliest availableeffective filing date and is a continuation application of U.S. patentapplication Ser. No. 14/771,975 titled “VANE ASSEMBLY FOR AN UNDUCTEDTHRUST PRODUCING SYSTEM” having a filing date of Sep. 1, 2015 and is aNational Phase of PCT/US2013/066403 filed on Oct. 23, 2013, and claimingpriority to provisional application No. 61/717,445 filed on Oct. 23,2012, and provisional application No. 61/771,451 filed on Oct. 23, 2012,and provisional application No. 61/771,314 filed on Mar. 1, 2013, andwhich are incorporated herein by reference in their entirety.

FIELD

The technology described herein relates to an unducted thrust producingsystem, particularly a vane assembly paired with rotating elements. Thetechnology is of particular benefit when applied to “open rotor” gasturbine engines.

BACKGROUND

Gas turbine engines employing an open rotor design architecture areknown. A turbofan engine operates on the principle that a central gasturbine core drives a bypass fan, the fan being located at a radiallocation between a nacelle of the engine and the engine core. An openrotor engine instead operates on the principle of having the bypass fanlocated outside of the engine nacelle. This permits the use of largerfan blades able to act upon a larger volume of air than for a turbofanengine, and thereby improves propulsive efficiency over conventionalengine designs.

Optimum performance has been found with an open rotor design having afan provided by two contra-rotating rotor assemblies, each rotorassembly carrying an array of airfoil blades located outside the enginenacelle. As used herein, “contra-rotational relationship” means that theblades of the first and second rotor assemblies are arranged to rotatein opposing directions to each other. Typically the blades of the firstand second rotor assemblies are arranged to rotate about a common axisin opposing directions, and are axially spaced apart along that axis.For example, the respective blades of the first rotor assembly andsecond rotor assembly may be co-axially mounted and spaced apart, withthe blades of the first rotor assembly configured to rotate clockwiseabout the axis and the blades of the second rotor assembly configured torotate counter-clockwise about the axis (or vice versa). In appearance,the fan blades of an open rotor engine resemble the propeller blades ofa conventional turboprop engine.

The use of contra-rotating rotor assemblies provides technicalchallenges. One such challenge is transmitting power from the powerturbine to drive the blades of the respective two rotor assemblies inopposing directions. A second challenge is minimizing the acousticsignature of the rotors. This is demanding because varied aircraftangles of attack cause the swirl angles into the rotor blades to varycircumferentially. The leading edges of blades with higher input swirlangles are loaded more heavily and tend to be more effective acousticradiators of the noise of the upstream rotor. Another challenge, in partrelated to minimizing acoustic signature of the rotors, arises withinstalling the rotors on an aircraft. Rotor blades located near aircraftflow surfaces, including, for example, wings, fuselages, and pylons, cancontribute to interaction penalties by disturbing the desireddistribution of flow seen by the aircraft flow surface. This leads tosuboptimal levels of resultant swirl into the wake of the aircraft andpropulsion system and reduced propulsive efficiency.

It would be desirable to provide an open rotor propulsion system whichmore efficiently integrates with an aircraft.

BRIEF DESCRIPTION

An unshrouded vane assembly for an unducted propulsion system includes aplurality of vanes which have non-uniform characteristics configured togenerate a desired vane exit swirl angle.

BRIEF DESCRIPTION OF THE DRAWINGS

The accompanying drawings, which are incorporated in and constitute apart of the specification, illustrate one or more embodiments and,together with the description, explain these embodiments. The drawingsinclude illustrations of radial sections taken through stages of axialflow airfoils and nearby aircraft surfaces, and are typically referredto as “roll-out-views.” These views are generated by sectioning airfoilstages and aircraft surfaces at a fixed radial dimension (measuredradially from the common airfoil stage centerline), then unrolling or‘rolling-out’ the sections to view them in two dimensional space whilemaintaining circumferential and axial relationships between the airfoilstages and aircraft surfaces. In all of the Figures which follow, likereference numerals are utilized to refer to like elements throughout thevarious embodiments depicted in the Figures. In the drawings:

FIG. 1 shows an elevational cross-sectional view of an exemplaryunducted thrust producing system;

FIG. 2 depicts graphically how various parameters such as camber andstagger angle are defined with respect to a blade or vane;

FIG. 3 shows a cross-sectional illustration “roll-out view” of anexemplary unducted thrust producing system with uniform vanes;

FIG. 4 shows a cross-sectional illustration “roll-out view” of anexemplary unducted thrust producing system with vanes with non-uniformstagger angle and non-uniform camber angle, as well as nearby aircraftsurfaces;

FIG. 5 shows a cross-sectional illustration “roll-out view” of anexemplary unducted thrust producing system with vanes with non-uniformstagger angle and non-uniform camber angle, with some vanes varying inaxial and circumferential position, as well as nearby aircraft surfaces;

FIG. 6 shows a cross-sectional illustration “roll-out view” of anexemplary unducted thrust producing system with vanes with non-uniformstagger angle and camber angle, with some non-uniform vanes in axial andcircumferential position, with vanes removed near aircraft surfaces;

FIG. 7 shows a cross-sectional illustration “roll-out view” of anexemplary unducted thrust producing system with vanes with non-uniformpitch angle;

FIG. 8 depicts an exemplary embodiment of a vane with pitch change viarigid body vane motion;

FIG. 9 is an illustration of an alternative embodiment of an exemplaryvane assembly for an unducted thrust producing system; and

FIG. 10 depicts vector diagrams illustrating Cu through both rows fortwo exemplary embodiments.

DETAILED DESCRIPTION

FIG. 1 shows an elevational cross-sectional view of an exemplaryunducted thrust producing system 70. As is seen from FIG. 1, theunducted thrust producing system 70 takes the form of an open rotorpropulsion system and has a rotating element in the form of rotatablepropeller assembly 20 on which is mounted an array of blades 21 around acentral longitudinal axis 80 of the propulsion system 70. Propulsionsystem 70 also includes in the exemplary embodiment a non-rotatingstationary element, vane assembly 30, which includes an array of vanes31 also disposed around central axis 80. For reference purposes, aforward direction for the unducted thrust producing system is depictedwith the arrow and reference letter F.

As shown in FIG. 1, the exemplary propulsion system 70 also includes adrive mechanism 40 which provides torque and power to the propellerassembly 20 through a transmission 50. In various embodiments, the drivemechanism 40 may be a gas turbine engine, an electric motor, an internalcombustion engine, or any other suitable source of torque and power andmay be located in proximity to the propeller assembly 20 or may beremotely located with a suitably configured transmission 50.Transmission 50 transfers power and torque from the drive mechanism 40to the propeller assembly 20 and may include one or more shafts,gearboxes, or other mechanical or fluid drive systems.

Blades 21 of propeller assembly 20 are sized, shaped, and configured toproduce thrust by moving a working fluid such as air in a direction Z asshown in FIG. 1 when the propeller assembly 20 is rotated in a givendirection around the longitudinal axis 80. In doing so, blades 21 imparta degree of swirl to the fluid as it travels in the direction Z. Vanes31 of the stationary element are sized, shaped, and configured todecrease the swirl magnitude of the fluid so as to increase the kineticenergy that generates thrust for a given shaft power input to therotating element. For both blades and vanes, span is defined as thedistance between root and tip. Vanes 31 may have a shorter span thanblades 21, as shown in FIG. 1, for example, 50% of the span of blades21, or may have longer span or the same span as blades 21 as desired.Dimension H in FIG. 1 represents the radial height of vane 31 measuredfrom longitudinal axis 80. Vanes 31 may be attached to an aircraftstructure associated with the propulsion system, as shown in FIG. 1, oranother aircraft structure such as a wing, pylon, or fuselage. Vanes 31of the stationary element may be fewer or greater in number than, or thesame in number as, the number of blades 21 of the rotating element andtypically greater than two, or greater than four, in number.

Vanes 31 may be positioned aerodynamically upstream of the blades 21 soas to serve as counter-swirl vanes, i.e., imparting tangential velocitywhich is opposite to the rotation direction of the propeller assembly20. Alternatively, and as shown in FIG. 1, vanes 31 may be positionedaerodynamically downstream of the blades 21 so as to serve as de-swirlvanes, i.e., imparting a change in tangential velocity which isgenerally counter to that of the propeller assembly 20.

FIG. 2 depicts graphically how various parameters such as camber andstagger angle are defined with respect to a blade or vane. An airfoilmeanline is a described as a line that bisects the airfoil thickness (oris equidistant from the suction surface and pressure surface) at alllocations. The meanline intersects the airfoil at leading edge andtrailing edge. The camber of an airfoil is defined as the angle changebetween the tangent to the airfoil meanline at the leading edge and thetangent to the angle meanline at the trailing edge. The stagger angle isdefined as the angle the chord line makes with the centreline axis.Reference line 44 is parallel to axis 11, and reference line 55 isorthogonal to reference line 44.

As mentioned above, FIG. 3 through FIG. 7 each include illustrations ofradial sections taken through stages of axial flow airfoils and nearbyaircraft surfaces, and are typically referred to as “roll-out-views.”These views are generated by sectioning airfoil stages and aircraftsurfaces at a fixed radial dimension measured radially from longitudinalaxis 80, reference dimension R in FIG. 1. When blades 21 and vanes 31 ofrespective propeller assembly 20 and vane assembly 30 are sectioned atreference dimension R, corresponding blade sections 22 and vanessections 32 are generated. Then the blade sections 22 and vanes sections32 are unrolled or ‘rolled-out’ to view the sections in two-dimensionalspace while maintaining the circumferential and axial relationshipsbetween the airfoil stages and any nearby aircraft surfaces. Referencedimension E for the axial spacing between blade sections 22 and vanesections 32. This allows the propeller assembly 20 and the vane assembly30 in FIG. 3 through FIG. 7 to be described in two dimensions. An axialdimension, parallel to the longitudinal axis 80 and generally alignedwith the direction Z of the moving working fluid shown in FIG. 1, and a‘rolled-out’ or flattened circumferential dimension X, orthogonal to theaxial dimension.

FIG. 3 describes a cross-sectional illustration “roll-out view” ofpropeller assembly 20 which as depicted has twelve blade sections 22.Each blade section 22 is individually labeled with lower case letters othrough z, with the section 22 labeled o repeating at the end of thesequence to highlight the actual circumferential nature of propellerassembly 20. Each blade section 22 has a blade leading edge 23. A linepositioned in the circumferential direction X through each blade leadingedge 23 defines a rotor plane 24. Each blade 21 and related section 22are spaced apart from each other and are located axially at the rotorplane 24.

Similar to the propeller assembly 20, the vane assembly 30 depicted inFIG. 3 has ten vanes sections 32, individually labeled a through j, eachwith a vane leading edge 33. A line positioned in the circumferentialdirection through each vane leading edge 33 defines a stator plane 34.In FIG. 3, each vane 31 and related section 32 in the vane assembly 30is identical in size, shape, and configuration, and is evenly spacedcircumferentially from each other, reference dimension P, and evenlyspaced axially from the rotor plane 24, reference dimension E. Anominal, evenly distributed circumferential spacing P, between vanes 31can be defined by the following equation using the radial height of thereference dimension R, and the number of vanes 31, N, in vane assembly30; P=R*2*π/N.

To optimize the installed performance and acoustic signature of thepropulsion system 70 when integrated with an aircraft, it may bedesirable to change the size, shape, configuration, axial spacingrelative to the rotor plane 24, and relative circumferential spacing ofeach vane 31 or group of vanes 31 and their related sections 32 in thevane assembly 30. Exemplary embodiments of this propeller system 20 andvane system 30 are shown in FIG. 4, FIG. 5 and FIG. 6. In each of thesefigures, the propeller assembly 20 and vane assembly 30 are locatedaxially forward of the aircraft surface 60. Additionally, an exemplaryembodiment of an aircraft surface 60 is represented as two wing sections61, and 62. Note that two wing sections are present in each “roll-outview,” because the radial section that generates these installed viewscuts through the wing of an aircraft in two circumferential locations.For the non-uniform vanes 31 in all of the Figures which follow, thisdashed and solid line depiction method is used to refer to exemplaryembodiments of nominal and non-nominal vane sections 32 respectively.

To minimize the acoustic signature it is again desirable to have theaerodynamic loading of the vane leading edges 32 to all be similar andbe generally not highly loaded. To maximize the efficiency and minimizethe acoustic signature of the propeller assembly 20, a desired goalwould be to minimize the variation in static pressure circumferentiallyalong the propeller assembly 20. To maximize the performance of the vaneassembly 30, another goal would be have to neither the aerodynamicloadings of the vane leading edges 32 nor the vane suction 35 andpressure surface 36 diffusion rates lead to separation of the flow.

To maximize the performance of the aircraft surface 60, depicted inthese exemplary embodiments as a wing sections 61 and 62, one goal maybe to keep the wing loading distribution as similar to the loadingdistribution the wing was designed for in isolation from the propulsionsystem 70, thus maintaining its desired design characteristics. The goalof maintaining the aircraft surface 60 performance as designed for inisolation from the propulsion system 70 applies for aircraft surfacesthat may be non-wing, including, for example, fuselages, pylons, and thelike. Furthermore, to maximize the performance of the overall aircraftand propulsion system 70 one of the goals would be to leave the lowestlevels of resultant swirl in the downstream wake. As described herein,the non-uniform characteristics of the vanes are tailored to accommodatethe effects of such an aircraft structure.

This optimal performance can be accomplished in part by developingnon-uniform vane exit flow angles, shown in FIG. 4 as angles Y and Z, tominimize interaction penalties of the installation and reduce acousticsignature. The first exemplary embodiment of this is shown in FIG. 4,where each vane 31 and related vane section 32 in the vane assembly 30are evenly spaced circumferentially from each other and evenly spacedaxially from the rotor plane 24. However, the nominal (without pitchchange) stagger angle and camber of the vane sections 32 in FIG. 4 varyto provide optimal exit flow angles into the aircraft surface 60,reference vane sections 32 labeled b through e, and g through i.

FIG. 5 shows another exemplary embodiment of vane assembly 30 providingflow complimentary to aircraft surface 60. In FIG. 5, vanes 31 andrelated vane sections 32 in vane assembly 30 are not evenly spacedcircumferentially from each other, nor are they evenly spaced axiallyfrom the rotor plane 24. The degree of non-uniformity may vary along thespan of a vane. Two vanes 31 are spaced axially forward of the statorplane 34, reference dimensions F and G, allowing the vane assembly 30 tomerge axially with the aircraft surface 60. The nominal (without pitchchange) stagger angle and camber angle of the vane sections 32 vary toprovide optimal exit flow angles into the wing sections 61 and 62, asshown in vane sections 32 labeled d through i.

FIG. 6 is similar to FIG. 5, but depicts the removal of two vanes 31adjacent to wing section 61. This exemplary embodiment allows the vanes31 to be evenly spaced axially from the rotor plane 24 and allows thewing section to merge axially with the vane assembly 30.

Although the location of the propeller system 20 and vane system 30 ineach of the foregoing exemplary embodiments was axially forward of theaircraft surface 60, it is foreseen that the propulsion system 70 couldbe located aft of the aircraft surface 60. In these instances, the priorenumerated goals for optimal installed performance are unchanged. It isdesirable that the propulsion system has suitable propeller assembly 20circumferential pressure variations, vane leading edge 32 aerodynamicloadings, and vane pressure surface 35 and suction surface 36 diffusionrates. This is accomplished in part by varying the size, shape, andconfiguration of each vane 31 and related vane section 32 in the vaneassembly 30 alone or in combination with changing the vane 31 pitchangles. For these embodiments, additional emphasis may be placed onassuring the combined propulsion system 70 and aircraft leave the lowestlevels of resultant swirl in the downstream wake.

The exemplary embodiment of the propeller assembly 20 and vane assembly30 in FIG. 3 is designed for a receiving a constant swirl angle,reference angle A, into vanes 31 along the stator plane 34. However, asthe aircraft angle of attack is varied the vanes move to off designconditions and the swirl angle into the vane assembly 30 will varyaround the stator plane 34. Therefore, to keep the aerodynamic loadingon the vane leading edges 33 roughly consistent along the stator plane34, a variable pitch system that would rotate either each vane 31 orgroup of vanes 31 a different amount is desirable. Such a pitch changecan be accomplished by rotating a vane 31 in a solid body rotation alongany axis, including, for example, the axis along the centroid of vanesection 32 or an axis along the vane leading edge 33. The desire forsimilar aerodynamic loading on the vane leading edges 33 is in partdriven by the desire to keep the acoustic signature of the propulsionsystem 70 low. Vanes 31 with high leading edge loadings tend to be moreeffective acoustic radiators of the noise created from the gust of theupstream propeller assembly 20. The exemplary embodiment of thepropeller system 20 and vane system 30 in FIG. 7 describes this desiredvariation in vane 31 via changes in vane section 32 pitch angles. Forease of explanation, we define the chord line angle of vanes at thedesign point as stagger and hence variations between vanes at the designpoint as stagger variations. As the engine moves to different operatingconditions, vanes may rotate around an axis referred to as pitch changeof the vanes. Variations in vane section chord angles that result fromthese sold body rotations are referred to as pitch angle variations.

In FIG. 7, each vane 31 and related vane section 32 in the vane assembly30 is identical in size, shape, and configuration, and are evenly spacedcircumferentially from each other and evenly spaced axially from therotor plane 24. However, the pitch angles of the vane sections 32 inFIG. 7 vary as they represent a change in the vane 31 pitch actuation toaccommodate varying input swirl, reference different input swirl anglesA and B, into stator plane 34 caused in part by changes in aircraftangle of attack. As desired, this provides similar aerodynamic loadingon the vane leading edges 33 to keep the acoustic signature of thepropulsion system 70 low. This similar loading can be accomplished byindependently changing pitch angle for individual blades or by changingpitch angles similarly for groups of vanes suitable for ganging. Thevanes 31 could rotate in pitch about any point in space, but it may bedesirable to maintain the original leading edge 33 circumferentialspacing and rotate the vanes 31 around a point at or near their leadingedge 32. This is shown in FIG. 7 using vane sections 32 labelled c, d,f, and g, where the nominal staggered vane sections 32 are depicted indashed lines and the rotated (or pitched) vane sections 32 are depictedas solid lines.

As shown by way of example in FIG. 8, it may be desirable that either orboth of the sets of blades 21 and vanes 31 incorporate a pitch changemechanism such that the blades and vanes can be rotated with respect toan axis of pitch rotation either independently or in conjunction withone another. Such pitch change can be utilized to vary thrust and/orswirl effects under various operating conditions, including to provide athrust reversing feature which may be useful in certain operatingconditions such as upon landing an aircraft.

The vane system 30, as suitable for a given variation of input swirl andaircraft surface 60 installation, has non-uniform characteristics orparameters of vanes with respect to one another selected either singlyor in combination from those which follow. A delta in stagger anglebetween neighboring vanes 31 and related vane sections 32 according toone embodiment of greater than or equal to about 2 degrees can beemployed, and according to another embodiment between about 3 degreesand about 20 degrees. A delta in camber angle between neighboring vanes31 and related vane sections 32 according to one embodiment of greaterthan or equal to about 2 degrees can be employed, and according toanother embodiment between about 3 degrees and about 15 degrees. Acircumferential spacing P at a given reference dimension R, betweenneighboring vanes 31 and related vane sections 32, for vane 31 counts Nfrom about 5 to about 30, from about 10% to about 400% of the nominal,even circumferential spacing can be employed. An axial spacing from therotor plane 24 to vanes 31 and related vane sections 32 up to about 400%of the radial height H, of the vane 31 can also be employed.

The non-uniform characteristic may be attributed to a portion of thespan of the vanes, or to substantially all of the span of the vanes.

The foregoing exemplary embodiments utilized twelve blades 21 and tenvanes 31, and one aircraft surface 60, but any combination of numbers ofblades 21, vanes 31, and aircraft surfaces 60 may be used.

It may be desirable to utilize the technologies described herein incombination with those described in commonly-assigned, co-pendingapplications [ ] and [ ].

In addition to configurations suited for use with a conventionalaircraft platform intended for horizontal flight, the technologydescribed herein could also be employed for helicopter and tilt rotorapplications and other lifting devices, as well as hovering devices.

The technology described herein is particularly beneficial for aircraftthat cruise with shaft power per unit annulus area of above 20 SHP/ft2(shaft horsepower per square foot) where the swirl losses can becomesignificant. Loadings of 20 SHP/ft2 and above permit aircraft to cruiseat Mach numbers above 0.6 Mach number without requiring excessivelylarge propeller areas to limit swirl losses. One of the major benefitsof the invention is its ability to achieve high shaft power per unitannulus area without significant swirl loss penalties and this opens theopportunity to cruise at Mach numbers of 0.8 and above.

Vanes 31 may optionally include an annular shroud or duct 100 distallyfrom axis 80 (as shown in FIG. 9) or may be unshrouded. In addition tothe noise reduction benefit the duct 100 provides a benefit forvibratory response and structural integrity of the stationary vanes 31by coupling them into an assembly forming an annular ring or one or morecircumferential sectors, i.e., segments forming portions of an annularring linking two or more vanes 31 such as pairs forming doublets. Theduct 100 may allow the pitch of the vanes to be varied as desired.

A significant, perhaps even dominant, portion of the noise generated bythe disclosed fan concept is associated with the interaction between thewakes and turbulent flow generated by the upstream blade-row and itsacceleration and impingement on the downstream blade-row surfaces. Byintroducing a partial duct acting as a shroud over the stationary vanes,the noise generated at the vane surface can be shielded to effectivelycreate a shadow zone in the far field thereby reducing overallannoyance. As the duct is increased in axial length, the efficiency ofacoustic radiation through the duct is further affected by thephenomenon of acoustic cut-off, which can be employed, as it is forconventional aircraft engines, to limit the sound radiating into thefar-field. Furthermore, the introduction of the shroud allows for theopportunity to integrate acoustic treatment as it is currently done forconventional aircraft engines to attenuate sound as it reflects orotherwise interacts with the liner. By introducing acoustically treatedsurfaces on both the interior side of the shroud and the hub surfacesupstream and downstream of the stationary vanes, multiple reflections ofacoustic waves emanating from the stationary vanes can be substantiallyattenuated.

FIG. 10 depicts the change in Cu across the rotating and stationaryelements, where Cu is the circumferential averaged tangential velocity.Vector diagrams are shown in a coordinate system in which the axialdirection is in the downward direction and tangential direction is leftto right. Multiplying the Cu times the airstream radius R gives theproperty RCu. The blade or vane loading at a given radius R is nowdefined as the change in RCu across the blade row (at a constant radiusor along a streamtube), here forth referred to as ARCu and is a measureof the elemental specific torque of said blade row. Desirably, the ARCufor the rotating element should be in the direction of rotationthroughout the span.

The foregoing description of the embodiments of the invention isprovided for illustrative purposes only and is not intended to limit thescope of the invention as defined in the appended claims. Othermodifications of the invention shall be apparent to those skilled in theart from the teachings herein, and it is, therefore, desired to besecured in the appended claims all such modifications as fall within thetrue spirit and scope of the invention.

What is claimed is:
 1. A thrust producing system, comprising: anaircraft structure comprising an aircraft surface; an unshroudedrotating element; a vane assembly located aft of the rotating element;wherein at least a portion of the aircraft surface is merged along anaxial direction with the vane assembly.
 2. The thrust producing systemof claim 1, wherein the aircraft structure comprises one or more of apylon, a fuselage, or a wing.
 3. The thrust producing system of claim 1,wherein a leading edge of the aircraft structure is merged along theaxial direction with the vane assembly.
 4. The thrust producing systemof claim 1, comprising: a drive mechanism configured to provide torqueand power to the unshrouded rotating element.
 5. The thrust producingsystem of claim 1, wherein the aircraft surface and the vane assemblyare together evenly spaced along the axial direction from a referencerotor plane.
 6. The thrust producing system of claim 1, wherein at leastthe vane assembly is configured to impart a change in tangentialvelocity of the air opposite to that imparted by the rotating element,and wherein the vane assembly comprises non-uniform characteristics withrespect to two or more vanes, and wherein the vane assembly isconfigured to generate a desired exit swirl angle.
 7. The thrustproducing system of claim 1, comprising: a plurality of vanes positionedat the vane assembly and the aircraft structure, wherein at least aportion of the plurality of vanes is variable in pitch.
 8. A thrustproducing system, comprising: an aircraft structure comprising anaircraft surface positioned at one or more of a pylon, a fuselage, or awing; an unshrouded rotating element; a drive mechanism configured toprovide torque and power to the unshrouded rotating element, the drivemechanism connected to an aircraft by the aircraft structure; anunshrouded vane assembly located aft of the rotating element; wherein atleast a portion of the aircraft surface is merged along an axialdirection with the vane assembly, and wherein at least the portion ofthe aircraft surface is positioned along a circumferential directionbetween two vanes of the vane assembly.
 9. The thrust producing systemof claim 8, wherein a leading edge of the aircraft structure is mergedalong the axial direction with the unshrouded vane assembly.
 10. Thethrust producing system of claim 8, wherein the aircraft surface and theunshrouded vane assembly are together evenly spaced along the axialdirection from a reference rotor plane.
 11. The thrust producing systemof claim 8, comprising: a non-rotating stationary element positionedalong the circumferential direction relative to a longitudinal axis ofthe thrust producing system, wherein the stationary element comprisesthe unshrouded vane assembly and the aircraft surface.
 12. The thrustproducing system of claim 11, wherein the stationary element isconfigured to impart a change in tangential velocity of the air oppositeto that imparted by the rotating element, and wherein the unshroudedvane assembly comprises non-uniform characteristics with respect to twoor more vanes, and wherein the stationary element is configured togenerate a desired exit swirl angle.
 13. The thrust producing system ofclaim 12, wherein the non-uniform characteristic is selected from thegroup consisting of: camber, stagger, circumferential spacing, axialposition, span, tip radius, and combinations thereof.
 14. The thrustproducing system of claim 8, comprising: a plurality of vanes positionedat the vane assembly and the aircraft surface.
 15. The thrust producingsystem of claim 14, wherein each of the plurality of vanes comprises aleading edge.
 16. The thrust producing system of claim 14, wherein atleast a portion of the plurality of vanes is variable in pitch.
 17. Athrust producing system for an aircraft, comprising: an aircraftstructure comprising a fuselage and a pylon, wherein the pylon comprisesa leading edge; an unshrouded rotating element; a drive mechanismconfigured to provide torque and power to the unshrouded rotatingelement, the drive mechanism connected to an aircraft by the aircraftstructure; an unshrouded vane assembly located aft of the rotatingelement and rotationally fixed in relation to a longitudinal axis of thedrive mechanism, wherein the unshrouded vane assembly comprises aplurality of vanes positioned along a circumferential direction; andwherein at least a portion of the leading edge of the pylon is mergedalong an axial direction between two vanes of the unshrouded vaneassembly.
 18. The thrust producing system of claim 17, wherein the pylonand the unshrouded vane assembly are together evenly spaced along theaxial direction from a reference rotor plane.
 19. The thrust producingsystem of claim 17, wherein at least the unshrouded vane assembly isconfigured to impart a change in tangential velocity of the air oppositeto that imparted by the unshrouded rotating element, and wherein theunshrouded vane assembly comprises non-uniform characteristics withrespect to two or more vanes, and wherein the unshrouded vane assemblyis configured to generate a desired exit swirl angle.
 20. The thrustproducing system of claim 17, wherein at least a portion of theplurality of vanes is variable in pitch.